Gas turbine bleed ecs cooling

ABSTRACT

A method for managing heat in an airplane includes extracting hot bleed air from a gas turbine engine, and cooling the hot bleed air in a first heat exchanger thereby forming warm bleed air. The method further includes expanding the warm bleed air with a turbine thereby forming cool bleed air, and using the cool bleed air as a heat sink in a second heat exchanger associated with an environmental control system and thereby forming used bleed air.

BACKGROUND

The present disclosure relates generally to airplanes and more specifically, to environmental control systems in airplanes.

Most modern military and commercial airplanes are powered by a gas turbine engine and include an Environmental Control System (ECS). Known ECS include air cooling and conditioning equipment such as compressors, valves, and heat exchangers, which are sometimes arranged into an ECS “pack”. Hot pressurized air is commonly taken from the gas turbine engine and ducted to the ECS pack for cooling and conditioning. This conditioned air is then provided to a cabin and/or flight deck for use by crew, passengers, and/or equipment on the airplane.

SUMMARY

A method for managing heat in an airplane is disclosed. The method includes extracting hot bleed air from a gas turbine engine, and cooling the hot bleed air in a first heat exchanger thereby forming warm bleed air. The method further includes expanding the warm bleed air with a turbine thereby forming cool bleed air, and using the cool bleed air as a heat sink in a second heat exchanger associated with an environmental control system thereby forming used bleed air.

A heat management system for an airplane is also disclosed. The system includes a first heat exchanger, a turbine, and a second heat exchanger. A gas turbine engine produces hot bleed air, and the first heat exchanger cools the hot bleed air thereby forming warm bleed air. The turbine expands the warm bleed air thereby forming cool bleed air, and the second heat exchanger uses the cool bleed air as a heat sink for a fluid in an environmental control system.

In another embodiment of the heat management system, the gas turbine engine has a bleed port for extracting hot bleed air. The first heat exchanger is fluidly connected to the bleed port and configured to cool the hot bleed air into warm bleed air. The turbine is fluidly connected to the first heat exchanger and configured to expand the warm bleed air into cool bleed air. The second heat exchanger is fluidly connected to the turbine and configured to cool a fluid for an environmental control system with the cool bleed air.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic of an airplane having a gas turbine engine.

FIG. 2 is a schematic of the gas turbine engine and a heat management system in accordance with the present disclosure.

FIG. 3 is a flow chart of a method for managing heat in accordance with the present disclosure.

DETAILED DESCRIPTION

As used herein, the term “airplane” includes any type of aircraft having an environmental control system. FIG. 1 is a perspective view of airplane 10 with cut-away to show a cross-section of gas turbine engine 12. Engine 12 includes compressor section 14, combustor section 16, and turbine section 18. Compressor section 14 further includes a fan 20, low pressure compressor (LPC) and high pressure compressor (HPC) 22. Primary air flow 24 travels through compressor section 14, combustor section 16, and turbine section 18 of engine 12 to generate power for airplane 10.

Engine 12 is attached to a wing of airplane 10 and includes compressor section 14, combustor section 16, and turbine section 18 in flow series. Fan 20, LPC and HPC 22 are located within compressor section 14 in flow series. In operation, primary air flow 24 enters Fan 20 of compressor section 14 and is compressed. Next, partially compressed primary air flow 24 enters LPC and HPC 22 of compressor section 14 and experiences further compression. Fully compressed primary air flow 24 then enters combustor section 16, which mixes fuel with fully compressed primary air flow 24 and combusts the mixture. Combusted primary air flow 24 enters turbine section 18 and experiences expansion, which forces one or more turbines to rotate. The rotating turbine then drives Fan 20, LPC and HPC 22.

FIG. 2 is a schematic of gas turbine engine 12 and heat management system 26 in accordance with the present disclosure. Engine 12 includes compressor section 14, combustor section 16, turbine section 18, fan 20, LPC and HPC 22, primary air flow 24, fan bleed port 28, low pressure or intermediate compressor (IC) bleed port 30, high pressure (HP) bleed port 32, and secondary air flow or hot bleed air 34 including fan bleed 34A, intermediate compressor (IC) bleed 34B, and high compressor (HC) bleed 34C. Heat management system 26 includes first heat exchanger 36, warm bleed air 38, turbine 40, cool bleed air 42, second heat exchanger 44, and used bleed air 46. Also depicted in FIG. 2 are first cool fluid 48, first warm fluid 50, and potential sources 52 for first cool fluid 48 including ram air 52A, lower stage (LS) bleed air 52B, or fuel 52C. Shaft 54 of turbine 40 is coupled to energy sink 56 including pump 56A, compressor 56B, generator 56C, or gearbox 56D. Second warm fluid 58, second cool fluid 60, and portions of environmental control system (ECS) 62 including vapor cycle system 62A, air cycle system 62B, and hybrid cooling system 62C are also depicted in FIG. 2. Hot bleed air 34 is extracted from engine 12, cooled, expanded, and used as a heat sink for a fluid in ECS 62.

As described above with reference to FIG. 1, engine 12 includes fan 20, LPC and HPC 22 of compressor section 14, combustor section 16, and turbine section 18 in flow series. Shown in FIG. 2 are three engine bleed ports: fan bleed port 28, IC bleed port 30, and HP bleed port 32. Fan bleed port 28 is located just downstream of fan 20 and upstream of LPC. IC bleed port 30 is located just downstream of LPC and upstream of HPC 22, downstream of fan bleed port 28 and upstream of combustor section 16. HP bleed port 32 is located at an end of HPC 22 just upstream of combustor section 16. Primary air flow 24 flows through fan 20, LPC and HPC 22 of compressor section 14, combustor section 16, and turbine section 18 of engine 12 to power airplane 10. A secondary air flow enters fan 20 along with primary air flow 24, but exits engine 12 at one or more of fan bleed port 28, IC bleed port 30, and HP bleed port 32. This secondary airflow or hot bleed air 34 can be extracted or “bled” from fan bleed port 28 as fan bleed air 34A, from IC bleed port 30 as IC bleed air 34B, and/or from HP bleed port 32 as HP bleed air 34C. One or more of fan bleed port 28, IC bleed port 30, and HP bleed port 32 are chosen depending on the type of airplane 10 and the application of hot bleed air 34. Once extracted from engine 12, hot bleed air 34 enters heat management system 26.

Heat management system 26 includes three subsystems; each subsystem is centered on one of first heat exchanger 36, turbine 40, and second heat exchanger 44. Each subsystem will be discussed in turn, beginning with first subsystem 26A of heat management system 26 centered upon first heat exchanger 36. Hot bleed air 34 enters an air inlet side of first heat exchanger 36 and warm bleed air 38 exits an air outlet side of first heat exchanger 36. First cool fluid 48 enters a fluid inlet side of first heat exchanger 36, and first warm fluid 50 exits a fluid outlet side of first heat exchanger 36. Source 52 of first cool fluid 48 can be ram air 52A, LS bleed air 52B, fuel 52C, or any other first cool fluid 48 that is cooler than hot bleed air 34. First cool fluid 48 is used as a heat sink for hot bleed air 34 within heat exchanger 36. First cool fluid 48 picks up heat from hot bleed air 34 such that hot bleed air 34 becomes warm bleed air 38 and first cool fluid 48 becomes first warm fluid 50.

After exiting first heat exchanger 36, warm bleed air 38 enters turbine 40, and second subsystem 26B_of heat management system 26. Warm bleed air 38 enters an air inlet of turbine 40 and cool bleed 42 air exits an air outlet of turbine 40. Shaft 54 attaches turbine 40 energy sink 56. Energy sink 56 can be a pump 56A, compressor 56B, generator 56C, gearbox 56D, or any other component requiring energy input. After being cooled by first heat exchanger 36, warm bleed air 38 is cooled further and expanded by turbine 40 to become cool bleed air 42. The work created by the expansion of warm bleed air 38 is transmitted through shaft 54 and used by energy sink 56. In other words, turbine 40 can provide energy to drive one or more of pump 56A, compressor 56B, generator 56C, or gearbox 56D.

After exiting turbine 40, cool bleed air 42 enters second heat exchanger 44, and third subsystem 26C of heat management system 26. Cool bleed air 42 enters an air inlet of second heat exchanger 44 and warm bleed air 46 exits an air outlet of second heat exchanger 44. Second warm fluid 58 enters a fluid inlet of second heat exchanger 44 and second cool fluid 60 exits a fluid outlet of second heat exchanger 44. Any portion of ECS 62 can provide second warm fluid 58 to second heat exchanger 44 including vapor cycle system 62A, air cycle system 62B, and/or hybrid cooling system 62C, so long as second warm fluid 58 is warmer than cool bleed air 42. Warm fluid 58 is used as a heat sink for cool bleed air 42 within second heat exchanger 44. Cool bleed air 42 picks up heat from second warm fluid 58 such that cold bleed air 42 becomes warm bleed air 46 and second warm fluid 58 becomes second cool fluid 60. In the depicted embodiment, second cool fluid 60 is returned to ECS 62 for use as a heat sink in one of vapor cycle system 62A, air cycle system 62B, or hybrid cooling system 62C. Accordingly, hot bleed air 34 from engine 12 is cooled and expanded into cool bleed air 42 and used as a heat sink for ECS 62.

FIG. 3 is a flow chart of the method for managing heat 64 in accordance with the present disclosure. Shown in FIG. 3 are steps of method 64: extract 66, cool 68, expand 70, heat exchange 72, and exhaust 74. Method 64 shows how to use hot bleed air 34 from engine 12 to cool a second warm fluid 58 from ECS 62. First, hot bleed air 34 is extracted from gas turbine engine 12 (extract step 66). As described with reference to FIG. 2, hot bleed air 34 can be extracted from one or more of fan bleed port 28, IP bleed port 30, or HP bleed port 32 as fan bleed air 34A, IP bleed air 34B, or HP bleed air 34C, respectively. Second, hot bleed air 34 is cooled in first heat exchanger 36 thereby forming warm bleed air 38 (cool step 68). As described with reference to FIG. 2, cool fluid 48 is obtained from any source 52 such as ram air 52A, LS bleed air 52B, fuel 52C. Third, warm bleed air 38 is expanded with turbine 40 thereby forming cool bleed air 42 (expand step 70). As described with reference to FIG. 2, turbine 40 is energetically attached to energy sink 56 such as pump 56A, compressor 56B, generator 56C, or gearbox 56D. Fourth, cool bleed air 42 is used as a heat sink in second heat exchanger 44 associated with ECS 62 (heat exchange step 72). Cool bleed air 42 becomes used bleed air 46 when used as a heat sink in second heat exchanger 44. Fifth, used bleed air 46 can be used as a heat sink for another portion of ECS 26 and/or exhausted outside of airplane (exhaust step 74).

Heat management system 26 and method for managing heat 64 are particularly advantageous for airplanes traveling at high mach speeds. At supersonic speeds, ambient/ram air and fuel may not be suitable for use as a heat sink because they may be simply too hot. Heat management system 26 and method for managing heat 64 provide an alternative to ambient/ram air and fuel whereby hot bleed air 34 from engine 12 is converted into cool bleed air 42 for use as a heat sink for a fluid in ECS 62.

While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims. 

1. A method for managing heat in an airplane, the method comprising: extracting hot bleed air from a gas turbine engine; cooling the hot bleed air in a first heat exchanger thereby forming warm bleed air; expanding the warm bleed air with a turbine thereby forming cool bleed air; and using the cool bleed air as a heat sink in a second heat exchanger associated with an environmental control system and thereby forming used bleed air.
 2. The method of claim 1, further comprising: exhausting the used bleed air out of the airplane.
 3. The method of claim 1, wherein the hot bleed air is extracted from one of a fan bleed port, an intermediate compressor bleed port, and a high compressor bleed port of the gas turbine engine.
 4. The method of claim 1, wherein the hot bleed air is cooled in the first heat exchanger by one of ram air, lower stage bleed air, or fuel.
 5. The method of claim 1, further comprising: operating one of a pump, a compressor, a generator, or a gearbox with energy generated by expanding the warm bleed air with the turbine.
 6. The method of claim 1, wherein the cool bleed air cools a working fluid in the second heat exchanger for one of an air cycle system, a vapor cycle system, or a hybrid cooling system.
 7. A heat management system for an airplane, the system comprising: a first heat exchanger for cooling the hot bleed air produced by a gas turbine engine, thereby forming warm bleed air; a turbine for expanding the warm bleed air thereby forming cool bleed air; and a second heat exchanger using the cool bleed air as a heat sink for a fluid in an environmental control system.
 8. The system of claim 7, wherein the hot bleed air is tapped from the gas turbine engine at one of a fan section, an intermediate compressor section, and a high compressor section.
 9. The system of claim 7, wherein one of ram air, lower stage bleed air, or fuel enter the first heat exchanger to provide the heat sink for cooling the hot bleed air.
 10. The system of claim 7, further comprising: a pump attached to the turbine, the pump for using work generated by the turbine expanding the warm bleed air.
 11. The system of claim 7, further comprising: a compressor attached to the turbine, the compressor for using work generated by the turbine expanding the warm bleed air.
 12. The system of claim 7, further comprising: a generator attached to the turbine, the generator for using work generated by the turbine expanding the warm bleed air.
 13. The system of claim 7, further comprising: a gearbox attached to the turbine, the gearbox for using work generated by the turbine expanding the warm bleed air.
 14. The system of claim 7, wherein the environmental control system is one of an air cycle system, a vapor cycle system, or a hybrid cooling system.
 15. A heat management system for an airplane, the system comprising: a gas turbine engine having a bleed port for extracting hot bleed air; a first heat exchanger fluidly connected to the bleed port, the heat first heat exchanger configured to cool the hot bleed air into warm bleed air; a turbine fluidly connected to the first heat exchanger, the turbine configured to expand the warm bleed air into cool bleed air; and a second heat exchanger fluidly connected to the turbine, the second heat exchanger configured to cool a fluid for an environmental control system with the cool bleed air.
 16. The system of claim 15, wherein the bleed port is located at one of a fan section, an intermediate compressor section, and a high compressor section.
 17. The system of claim 15, wherein one of ram air, lower stage bleed air, or fuel enter the first heat exchanger to cool the hot bleed air into warm bleed air.
 18. The system of claim 15, wherein one of a pump, a compressor, a generator, or a gearbox is attached to the turbine for using work generated by expanding the warm bleed air.
 19. The system of claim 15, wherein the environmental control system is one of an air cycle system, a vapor cycle system, or a hybrid cooling system. 